The invention relates to a variable cycle gas turbine engine of the bypass type and, more particularly, to a variable cycle gas turbine engine suitable for powering a supersonic aircraft wherein the engine bypass ratio and gas flow may be controlled to satisfy particular engine operating conditions.
Considerable attention has been devoted to developing a gas turbine engine with the high specific thrust characteristics of a turbojet or low bypass turbofan at supersonic speeds which can also be configured to exhibit the lower specific thrust, low noise and low fuel consumption characteristic of a high bypass turbofan at subsonic speeds in order that a mixed mission aircraft may be developed.
To this end, modern aircraft designers have strived to develop the aircraft engine design criteria which would enable the development of a suitable mixed mission aircraft. Several design approaches to this problem have been offered. However, all such prior art approaches have failed to result in an engine with sufficient flow flexibility to enable efficient, stall free operation in all modes. Such prior art systems have included various concepts of retractable fans, variable area turbines, variable pitch fans, as well as more exotic and highly complex techniques such as those utilizing combinations of turbofan and turbojet engines in tandem or concentric flow relation. In addition to a lack of flow flexibility, these more exotic arrangements have the obvious disadvantage of being highly inefficient due to the dead weight associated with those engine components not used in all modes of flight.
More recent attempts at developing practical variable cycle engines include the selective direction of the inlet fan stream through alternative upstream fan ducts using inverter valves. While more effective than prior attempts at achieving satisfactory mixed mission performance, such systems have exhibited several negative characteristics. These include the addition of extra undesired length, weight and complexity to the engine.
Another such prior art system is disclosed in U.S. Pat. No. 3,635,029 issued to Claude Charles Felox Menioux on Jan. 18, 1972. In the Menioux system, a gas turbine engine of the duct barrier type is configured to operate as a ramjet or as a turbofan engine by means of a valve downstream of the core engine. One major disadvantage of this type is that the outer duct must be designed to accommodate extremely high temperatures since it must handle the high temperature core gas stream as well as the lower temperature bypass gas stream. In addition, the presence of the burners in the outer duct create still other high temperature problems in the design of the outer duct.
A further disadvantage of the gas turbine engine disclosed in the Menioux patent and other prior art variable cycle engines is that they fail to meet desired performance goals in all modes of operation because they have insufficient flow variability to maintain satisfactory engine performance in both supersonic and subsonic flight.
One of the reasons that prior art fixed cycle and variable cycle engines have not maintained satisfactory performance at both supersonic and subsonic speed is that the airflow to the inlet of such engines in not matched to the inlet airflow potential during all phases of flight. Typically in such engines the inlet is sized to be full at the maximum thrust of the engine. However, as engine thrust is decreased below the cruise thrust, typically by decreasing the bypass ratio, the engine airflow demand is considerably less than the total airflow supplied to the inlet. This excess of airflow causes inlet spillage drag which significantly increases the installed fuel consumption of prior art variable cycle engines.